1. Field of the Invention
The invention is generally related to rockets and more particularly to solar thermal rockets.
2. General Background
Solar thermal rockets were first proposed in 1954 as a way to provide greater specific impulse than chemical rockets. Solar thermal rockets use the sun's energy to heat a propellant (typically hydrogen) to extremely high temperatures and then expel the hot gas through a nozzle to provide thrust. The high temperature and low molecular weight of the propellant combine to produce a specific impulse of two to four times that of a chemical rocket. Generally, solar thermal rockets have been of either a "direct gain" design in which the propellant is heated directly by very large solar collectors during a long continuous burn, or of a "thermal energy storage" design which collects and stores energy from smaller collectors for use in short impulsive "burns". Recently, a "Hybrid Direct Gain/Thermal Energy Storage" design has been proposed that combines the high-temperature propellant capability of the direct gain design with the smaller collector feature of the thermal energy storage design. Each of these designs has advantages and disadvantages.
The direct gain rocket requires very large solar collectors (concentrators) to provide sufficient energy to heat the hydrogen propellant as it passes through a cavity comprised of refractory metal tubes or encapsulated foam (typically rhenium). The advantage of this type of rocket is that the temperature of the propellant can be extremely high (theoretically greater than 3,000 K), thus providing high specific impulse thrust. The drawback with this design is that the solar collector(s) must be extremely large (often twenty-five to fifty meters in diameter) to provide the energy needed to heat the propellant from its stored cryogenic state (25 K) to the very high thrust temperature. Concentrator technology has not matured to the point where such concentrators are available for space applications (i.e. light weight and small stowed volume that fit existing space launch vehicles).
The thermal energy storage design solves the concentrator problem by collecting and storing solar energy over an orbital period, and then using the stored energy to provide thrust for a short impulsive burn. A number of burns are required to get the spacecraft to its destination. The longer the storage phase of the mission, the smaller the collector can be. This approach allows the use of existing collector technology to enable the development of a rocket. However, the major drawback to such a system is that the energy storage materials (typically rhenium coated graphite or tungsten encapsulated boron nitride) have temperature limitations well below that of a direct gain system. Current storage designs are limited to about 2400 K to avoid excessive carbon diffusion through the rhenium cladding. Thermal shock, which occurs when the hot thermal storage material/cladding is initially subjected to high velocity cold propellant, can also be a problem in thermal energy storage designs. Another problem is that the temperature of the heated propellant is very high at the start of the pulse but after a short period decreases as heat is extracted by the cold propellant. The resultant performance is less than that theoretically possible using the direct gain design with extremely high propellant outlet temperatures.
The Hybrid Direct Gain/Thermal Energy Storage design adds an all refractory metal section following the thermal energy storage section to allow heating of the propellant above the temperature limit of the thermal energy storage materials. The higher temperatures improve orbit transfer performance.
A problem with both the direct gain and the hybrid designs is that retention of thermal energy becomes much more difficult as the peak cavity temperature increases. Multi-foil insulation is often used to confine heat to the hot zone. At very high temperatures, heat loss out of the cavity aperture and through the multi-foil insulation is dominated by radiation heat transfer, which varies with temperature to the fourth power. Raising the cavity temperature by ten percent results in more than a forty-percent increase in heat loss. Heat input must be significantly increased to compensate for the larger heat losses if very high temperatures are to be obtained. As previously indicated, the size of the space deployable solar concentrators is already a limiting factor in solar thermal rocket systems.
Some solar powered rocket systems incorporate a secondary concentrator between the primary concentrator and the cavity to reduce the size of the aperture, which in turn reduces the amount of heat that can escape by radiation out of the cavity. Losses due to less-than-perfect reflection in the secondary concentrator can be significant. Although it is theoretically possible to limit these losses to less than ten percent, realistic losses are on the order of fifteen to thirty-five percent of the power from the primary concentrator.
In addition to high heat losses, very high temperature designs present material, structural, and manufacturing challenges that are difficult to solve individually and even more difficult to solve in a system, particularly under normal budget constraints. At the desired very high temperatures, refractory materials must be used. Rhenium is the preferred material because of its compatibility with hydrogen and carbon, ductile behavior over the entire temperature range, low vapor pressure, high strength, and high modulus of elasticity. Unfortunately, it is expensive, difficult to form and join, very dense, and has a low heat capacity. The structural behavior of rhenium at very high temperatures is not well characterized and varies significantly with only slight variations in manufacturing processes. To effectively capture and transfer heat to the propellant will require fabrication of leak tight components with relatively large surface areas. Although rhenium-processing technology is advancing, experience with making reliable, leak tight, efficient and lightweight rhenium heat exchangers has proven to be difficult.
As discussed, the above concepts offer enhanced performance, but each has difficult engineering problems particularly when associated with extremely high temperature. A more practical approach is needed.